A gas turbine engine comprises a turbine including one or more stages of turbine vanes and one or more stages of turbine blades. Each stage of turbine vanes comprises a plurality of circumferentially spaced turbine vanes and each stage of turbine blades comprises a plurality of circumferentially spaced turbine blades. The stages of turbine vanes and turbine blades are arranged alternately in flow series. The turbine vanes are mounted on static structures of the gas turbine engine whereas the turbine blades are mounted on rotatable structures of the gas turbine engine. A turbine vane generally comprises an aerofoil and two platforms and the aerofoil extends between and is secured to the platforms whereas a turbine blade generally comprises an aerofoil, a platform and a root and the aerofoil and root are secured to and extend in opposite directions from the platform.
The turbine vanes and turbine blades are located in a position downstream of a combustion chamber of the gas turbine engine and are exposed to gases at very high temperatures. In order to enable the turbine vanes and turbine blades to operate at these high temperatures the turbine vanes and turbine blades are manufactured from superalloys, are impingement and/or film cooled and are provided with thermal barrier coatings.
The platforms of the turbine vanes require a considerable amount of thermal protection whereas the aerofoils of the turbine vanes require a lesser amount of thermal protection. The thermal protection of the platforms of the turbine vanes comprises a metallic bond coating and a ceramic thermal barrier coating. The metallic bond coating is deposited onto the platforms by plasma spraying and the ceramic thermal barrier coating is deposited onto the metallic bond coating on the platforms by plasma spraying. The ceramic thermal barrier coating may have a thickness of up to 1 mm. Without these coatings on the platforms of the turbine vanes the platforms would be burnt away quickly resulting in a short service, operating, life for the turbine vanes. The aerofoils of the turbine vanes rely on film cooling to control the temperature of the aerofoil in conjunction with thermal protection. The thermal protection of the aerofoils of the turbine vanes also comprises a metallic bond coating and a ceramic thermal barrier coating. The metallic bond coating is deposited onto the aerofoils by plasma spraying and the ceramic thermal barrier coating is deposited onto the metallic bond coating on the aerofoils by physical vapour deposition (PVD), e.g. electron beam physical vapour deposition. The ceramic thermal barrier coating may have a thickness of up to 0.12 mm.
The turbine vanes have a fillet radius, a curved transition, from the aerofoil to the respective platform and the fillet radius is designed to optimise the aerodynamic performance of the turbine vanes.
However, the difference in the thickness of the ceramic thermal barrier coatings mentioned above on the platforms and the aerofoil of a turbine vane results in a step at the junction between each platform and the aerofoil rather than a smooth fillet radius, smooth curved transition, from the platform to the aerofoil. The step at each junction between the aerofoil and the platform of the turbine vane produces undesirable aerodynamic losses and thermodynamic characteristics. The aerodynamic losses from the steps at the junctions between the aerofoils and the platforms of the turbine vanes produce performance losses in the turbine and the gas turbine engine as a whole which results in increased fuel consumption.
The present disclosure seeks to provide a method of coating a turbine vane and a coated turbine vane which reduces or overcomes this problem.